Daniele Bianchi

Thermochemical erosion analysis for graphite/carbon-carbon rocket nozzles

A study is conducted to predict graphite/carbon–carbon nozzle erosion behavior in solid rocket motors for wide variations of propellant formulations. The numerical model considers the solution of Reynolds-averaged Navier– Stokes equations in the nozzle, heterogeneous chemical reactions at the nozzle surface, variable transport and thermodynamic properties, andheat conduction in the nozzlematerial. Twodifferent ablationmodels are considered and compared: a surface equilibriumapproach and afinite-ratemodel. Results show that the erosion rate is diffusion limited for metallized propellants, ensuring sufficiently high wall temperatures, and it is kinetic limited for nonmetallized propellants. For low surface temperatures, the twomodels are consistent with each other and predict the same erosion rate, while the surface equilibrium model overpredicts the recession at low surface temperatures. The calculated results show an excellent agreement with the experimental data from the ballistic test and evaluation system motor firings, and the finite-rate model actually improves the predictions when the kinetic-limited regime is approached.

Coupled Analysis of Flow and Surface Ablation in Carbon-Carbon Rocket Nozzles

A study is conducted to predict C/C nozzle recession behavior in solid rocket motors for broad variations of propellant formulations and motor operating conditions. The numerical model considers the turbulent flow in the nozzle, heterogeneous chemical reactions at the nozzle surface, variable transport and thermodynamic properties, and heat conduction in the nozzle material. Results show that the recession rate is largely determined by the diffusion of the major oxidizing species (H2O, CO2, OH) to the nozzle surface. Both the concentration of the major oxidizing species -affected by the aluminum content of the propellant- and the chamber pressure exert a strong influence on the recession rate. The erosion rate increases almost linearly with chamber pressure and decreases with propellants with higher aluminum content. The calculated results show a very good agreement with the experimental data from the BATES motor firings.

Navier-Stokes Simulations of Hypersonic Flows with Coupled Graphite Ablation

equilibriumablationwithsurfacemassandenergybalancesfullycoupledwiththenumericalsolverandcanaccount for both surface oxidation and sublimation. The surface temperature is obtained from the steady-state ablation approximation. This numerical procedure can predict aerothermal heating, chemical species concentrations, and carbon material ablation rate over the heat-shield surface of reentry vehicles. Two-dimensional axisymmetric simulations have been performed to numerically reproduce the ablation of a graphite sphere cone that has been tested in the Interaction Heating Facility at the NASA Ames Research Center. The freestream conditions of the selected test case are typical for Earth reentry from a planetary mission. The predicted ablation rate and surface temperatureassumingfrozenchemistryinthe flowshowagoodagreementwiththeavailableexperimentaldata.The agreementisfurtherimprovedfreezingthenitrogenrecombinationreactionatthesurfacetobemoreconsistentwith experimental observation, which has shown nitrogen atom recombination not to occur at the graphite surface.

Chemical Erosion of Carbon-Phenolic Rocket Nozzles with Finite-Rate Surface Chemistry

Ablative materials are commonly used to protect the nozzle metallic housing and to provide the internal contour to expand the exhaust gases in solid rocket motors. Because of the extremely harsh environment in which these materials operate, they are eroded during motor firing with a resulting nominal performance reduction. The objective of the present work is to study the thermochemical erosion behavior of carbon-phenolic material in solid rocket motor nozzles. The adopted approach relies on a validated full Navier–Stokes flow solver coupled with a thermochemical ablation model, which takes into account finite-rate heterogeneous chemical reactions at the nozzle surface, rate of diffusion of the species through the boundary layer, pyrolysis gas and char-oxidation product species injection in the boundary layer, heat conduction inside the nozzle material, and variable multispecies thermophysical properties. The results obtained with the proposed approach are compared with two sets of experimental data: subs...

Simulation of Gaseous Oxygen/Hydroxyl-Terminated Polybutadiene Hybrid Rocket Flowfields and Comparison with Experiments

Numerical simulations of the flowfield in a gaseous oxygen/hydroxyl-terminated polybutadiene hybrid rocket engine are carried out with a Reynolds-averaged Navier–Stokes solver including detailed gas/surface interaction modeling based on surface mass and energy balances. Fuel pyrolysis is modeled via finite-rate Arrhenius kinetics. A simplified two-step global reaction mechanism is considered for the gas-phase chemistry to model the combustion of 1,3-butadiene in oxygen. Results are compared with the firing test data obtained from a laboratory-scale hybrid rocket in which gaseous oxygen is fed into axisymmetric hydroxyl-terminated polybutadiene cylindrical grains through an axial conical subsonic nozzle. With the oxidizer fed by this injector, which generates nonuniform conditions at the entrance of the fuel port, the fuel regression rate is shown to increase several times with respect to the case of homogeneous injection of the oxidizer through all the grain port area, in agreement with the experimental f...

Carbon-Carbon Nozzle Erosion and Shape Change in Full-Scale Solid-Rocket Motors

The erosion of nozzle protection materials during solid-rocket-motor burning needs to be accounted for to get reliable performance predictions, especially for long-durationfirings.A study is conducted topredict carbon–carbon nozzle erosion behavior in full-scale solid-rocket motors for wide variations of motor operating conditions. The numerical model considers the solution of Reynolds-averaged Navier–Stokes equations in the nozzle, heterogeneous chemical reactions at the nozzle surface, ablation species injection in the boundary layer, variable multicomponent transport and thermodynamic properties, and heat conduction in the nozzlematerial. Two different ablationmodels are considered: a diffusion-limited approach and a finite-rate approach. The numerical model is used to study the erosion of carbon–carbon nozzle inserts for the secondand third-stage solid-rocket motors of the European Vega launcher. The effect of variable chamber pressure over the burning time and the effect of nozzle shape change on the erosion rate are taken into account in the numerical analysis. The obtained results show a very good agreement with the measured final eroded profile along the entire carbon–carbon nozzle throat insert for both motors. The shapechange effect is shown to be an important factor that has to be taken into account to get a goodprediction of the throat erosion for long-duration firings.

A numerical approach for high-temperatures flows over ablating surfaces

CFD codes typically treat fluid/solid boundary conditions in a simplified manner such as constant prescribed temperature or heat flux with zero mass transfer. However, thermal protection materials (TPS) strongly interact with the flow so that simple CFD surface boundary conditions cannot realistically be used for TPS design. In order to obtain a better estimation of the wall heat flux over an ablating surface, a two-dimensional axisymmetric full Navier-Stokes equation solver is used coupled with surface mass and energy balance and an equilibrium ablation model for graphite. Flat plate tests are presented. Results are compared with the most commonly used simplified approaches.

Radiation and Roughness Effects on Nozzle Thermochemical Erosion in Solid Rocket Motors

Surface roughness and radiation effects on the erosion behavior of a graphite nozzle are studied for both metallized and nonmetallized propellants. A validated numerical approach that relies on a full Navier–Stokes flow solver coupled with a thermochemical ablation model is used for the analysis. A modification of the Spalart–Allmaras turbulence model is implemented to account for surface roughness. Net radiative heat flux is considered in the surface energy balance at the nozzle interface. Two different simplified models are used to evaluate the integral emissivity of dispersed alumina particles. Individual and combined effects of roughness and radiation are analyzed. Surface roughness enhances the erosion rate for both metallized and nonmetallized propellants noticeably. The radiation influences the erosion rate of nonmetallized propellant more than the metallized one, mainly due to the different erosion regimes, kinetically controlled for the former and diffusion controlled for the latter.

Thermochemical Erosion Analysis of Carbon-Carbon Nozzles in Solid-Propellant Rocket Motors

The erosion of rocket-nozzle materials during motor firing is one of the major limits in the advancement of solid-rocket propulsion. A study is conducted to predict carboncarbon nozzle erosion behavior in solid rocket motors for wide variations of motor operating conditions. The numerical model considers the solution of Reynolds averaged NavierStokes equations in the nozzle, heterogeneous chemical reactions at the nozzle surface, variable multi-component transport and thermodynamic properties, and heat conduction in the nozzle material. Two different ablation models are considered: a surface equilibrium approach and a finite-rate approach. The numerical model is used to study the carboncarbon nozzle throat insert erosion of the European VEGA launcher third stage and second stage solid rocket motors. The effect of variable chamber pressure over the burning time and the effect of nozzle shape change on the erosion rate are taken into account in the numerical analysis. The obtained results show a very good agreement with the measured final eroded profile along the entire carbon-carbon nozzle throat insert for both motors.

Chemical reaction effects on heat loads of CH4/O2 and H2/O2 rockets

Cooling of liquid rocket thrust chamber walls to allowable solid material temperatures induces near-wall chemical reactions, which are known to have an important role on the heat transfer from the hot gas to the wall. In this study, the contribution of near-wall chemical reactions to heat flux is investigated and quantified by suitable numerical analyses. Numerical results are first compared to literature experimental data of wall heat flux in subscale calorimetric thrust chambers for both oxygen/methane and oxygen/hydrogen propellant combination. Then, a parametric analysis is carried out varying chamber pressure, wall temperature, and propellant combination. This study highlights that oxygen/methane combustion products are more subject to near-wall recombination phenomena. They provide an increase of wall heat flux between 20 and 30% with respect to the frozen flow model evaluation, whereas in the case of oxygen/hydrogen, the wall heat flux increase due to recombination reactions is between 7 and 14%.